It will be understood that gas turbine engines are utilised with regard to aircraft in order to provide propulsion in order to drive the aircraft both on the ground as well as provide propulsion in the air. In such circumstances the power required will vary dependent upon the operational phase of the aircraft. It will also be understood that these gas turbine engines also provide power in order to drive electrical generators and other auxiliary devices such as hydraulic or fuel pumps for the aircraft and the engine itself. These generators and auxiliary devices are typically driven through a gear box drive coupling to a shaft of the engine.
Increasing use of electrical actuators and devices to provide auxiliary functions as well as greater use of electronic processor and other control devices can significantly increase the power generation requirements from electrical generators driven by the primary engine of an aircraft. It will also be understood that the power demands of electrical generators to supply sufficient electrical power to the control and other auxiliary devices can significantly vary between aircraft phases such as ground operations and in-flight cruising. Such variations significantly complicate engine control.
As indicated above typically electrical generators as well as other auxiliary devices will be driven through a gearbox coupling to a shaft within an engine. Normally, an engine arrangement will incorporate several shafts and therefore when power is taken from one shaft to drive electrical generators/machines and other auxiliary devices imbalances may occur within the engine.
A particular compressor geometry operating in a stable condition will have a unique relationship between the non dimensional flow and the operating pressure ratio between the rear and front of the compressor. This is called the working line. If the compressor is driven to operate at the same flow but a higher pressure ratio it will eventually surge. There is also a unique relationship between non dimensional flow and the pressure ratio for surge. This is called the surge line.
Handling bleed valves are used to control the operation of axial flow compressors. They are located either at the back of a compressor in which case they affect the working line of that compressor or they are located mid way through the compressor in which case they affect the working line and the surge (also known as stall) line of the compressor.
The traditional method of control of a handling bleed valve is by a bleed schedule. Engine parameters which can easily be measured such as shaft speed and temperature are used to calculate the compressor non dimensional flow giving a measure of where the compressor is operating in terms of surge etc. Because the available margin between the working line and surge line varies with flow it is possible to schedule bleed valves to open at certain flows and closed at other flows. The schedule is set up to maintain sufficient margin between the surge and working lines to account for all the variation in surge and working lines possible during service life of the engine.
During transient operation the compressor will move away from the steady state working line due to a variety of thermodynamic effects. In this case more margin is required and a separate transient bleed schedule is required. Further advances include separate bleed schedules for Approach Idle, Reverse thrust operation, detection of water ingestion and detection of surge. Each schedule is designed to alter the engine matching in the most advantageous way.
It is known to control the surge margin of a compressor based on thermodynamic changes that occur within that compressor. No account is taken of engine matching changes which occur outside that compressor. Recent requirements for the aerospace gas turbines to provide very large levels of core power to the aircraft electrical generators can cause problems. Typically, the power requirements have gone from around 300 hp to over 800 hp. In addition these power requirements must be able to be provided at any time by the gas turbine with no prior knowledge or warning.
The power extraction is taken using one or more generators that are attached directly to the engine by means of a gearing system which is connected to one of the shafts. A drag force caused by the generators generating electricity is applied directly to the shaft in question. This means that some proportion of the power from the turbine attached to that shaft is used in providing power off take rather than driving the compressor on that shaft. The result is that for a given set of compressor inlet conditions the pressure ratio driven by the compressor will be lower i.e. the working line will drop.
For multi shaft engines being controlled to a constant thrust the drop in the working line of one of the compressors will result in more fuel being introduced and a consequent increase in the speed of the unloaded shaft or shafts. The precise rematching of the engine will depend on the exact engine configuration and the environmental conditions in terms of altitude and mach number.
Large power off take causes particular problems at low engine powers and using prior engine bleed control methods would force an unacceptably high minimum idle level or alternatively limit the practical level of power off take to a level significantly below desired requirements. Even a slightly raised idle level would be unacceptable because it would impact the aircraft's ability to manoeuvre during descent, would cause brake wear and other safety issues during taxi operations and would give a very high mission fuel burn penalty.